The wing or empennage structure of a modem airliner is typically a stiffened skin construction. Together with spars and ribs, the skin forms a torque box which will resist external loads.
In the case of a fuselage, the curved stiffened skin panels together with fuselage frames form the stiffened shell. Traditionally these skins are made from aluminium alloys, but as aircraft performance is becoming more and more important, composite skin panels are becoming more and more popular in aircraft primary structure construction.
Stiffened composite panels in primary structures may be used in horizontal tail plane, vertical tail plane and/or centre wing box construction. Typical for all these (excluding centre wing box) is that the skin is manufactured starting from the aerodynamic or outer surface. Since the thickness tolerance of the components is relatively poor this leads to additional costs in the final assembly where the resulting gaps and/or mismatches between skins, ribs and spars must be filled or adjusted with a suitable method to maintain the aerodynamic tolerance of the whole torque box.
Significant savings in the final assembly phase and completely new torque box designs could be utilised if the skin thickness tolerance could be maintained accurately enough so that both the outer mould line (aerodynamic) and inner mould line (e.g. spar & rib landings, main landing gear area) tolerances in critical locations could be controlled.
U.S. Pat. No. 4,683,018 describes a method of composite material manufacturing process in which a stack is placed on an upwardly directed female former, and then stamped by a male hydraulic press.